Method of surface charge distribution

ABSTRACT

A lightning strike protection system for protecting composite structures, an improved lightning strike appliqué (LSA) for such a lightning strike protection system, and a method of protecting composite structures, such as an aircraft fuselage. The LSA is electrically connected to adjacent conductive surfaces, e.g., by a fuzz button or a wire bond inserted in the bottom of the LSA. An adjacent conductive surface may be another LSA, a lightning diverter overlay, or a current return network. Charge, e.g., from a lightning strike to the LSA, flows to the conductive layer through the electrical connector.

CROSS REFERENCE TO RELATED APPLICATION

The present application is a continuation of allowed U.S. patentapplication Ser. No. 11/615,786, Publication No. 2010/0134945,“Electrical Connects For Charge Distribution Appliqué,” to Diane C.Rawlings et al., filed Dec. 22, 2006; a continuation in part of U.S.Pat. No. 7,525,785, “Lightning Strike Protection Method and Apparatus,”to Diane C. Rawlings, filed Dec. 14, 2006 and issued Apr. 28, 2009; andis related to Published U.S. patent application Ser. No. 10/941,429,Publication No. 2005/0181163, “Appliqué,” filed Sep. 15, 2004 andpublished Aug. 18, 2005; and to allowed U.S. patent application Ser. No.11/229,911, Publication No. 2006/051592, “Wide Area Lightning DiverterOverlay,” filed Sep. 19, 2005 and published Mar. 9, 2006, both to DianeC. Rawlings et al., all assigned to the assignee of the presentapplication.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention generally relates to protecting compositestructures from lightning strikes, and more particularly, electricallyconnecting multiple conductive appliqués together in a continuouscurrent path, for example, for establishing a preferred current path toground along the areas protected by Lightning Strike Appliqué, orelectrically connecting the appliqué to a ground contact for thepurposes of preventing the buildup precipitation-static (p-static)charge on the aircraft surface or on underlying components.

2. Background Description

The ability to effectively manage lightning strikes on compositematerials that form structural panels for wings, fuselages, fuel tanks,and other components of an aircraft structure is an importantconsideration for the safety of an aircraft.

Composite materials are highly desirable for use as structuralcomponents due to their lower mass, while possessing excellentstructural rigidity and high strength. However, composite materials arenot highly conductive and cannot dissipate the energy from a lightningstrike as efficiently as traditional metal body components used in manyconventional aircraft.

Carbon fiber reinforced plastic (CFRP) is one type of composite materialused for skin, spar and rib installations on aircraft. A CFRP structureis about 2000 times more resistive than most metals, and consequentlyCFRP is more prone to electrical breakdown when subjected to currentsfrom lightning strikes, especially at interfaces and fasteners.

Moreover, protection is needed against lightning strikes for not onlycomposite skins and underlying structures, but for sensitive equipment,like hydraulic lines and fuel tanks, as well.

Appliqué coatings, such as Lightning Strike Appliqué (LSA), whichcontain a thin metal foil, and Wide Area Lightning Diverter Overlay(WALDO), are used to protect aircraft. These coatings are described indetail in US Patent Application 2006/0051592, which is incorporatedherein by reference.

When using a lightning protection approach, such as LSA/WALDO, toprotect the composite structure it is important to reliably transitionthe current that is carried by or on the appliqué coating system to agrounded metallic structural component or current return network.

Typical current return networks used on aircraft are buried inside thestructure. This solution forces a designer to drive high electricalcurrents into the skin and composite structure itself. High currentsdamage sites at each electrical discontinuity, including fasteners,joints, fiber interfaces, panel edges, and the like, as well as creatinghot spots, edge-glow or sparks, which, for example, could ignite thefuel within the wing box.

The difficulty of predicting where currents go once an aircraft isstruck by lightning, leads to over-designing many areas of the structureand to the duplication of protection schemes.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing and other objects, aspects and advantages will be betterunderstood from the following detailed description of a preferredembodiment of the invention with reference to the drawings, in which:

FIGS. 1A-B show an example of an aircraft protected by conductor withlightning strike appliqué (LSA) gores electrically attached andconnected through a flexible, electrical connection, according to anembodiment of the present invention.

FIG. 2A shows a cross sectional example, wherein conductor is a LSA goreoverlapped by another LSA gore with both adhesively attached tocomposite skin;

FIG. 2B shows a cross section of connection of a LSA gore to anunderlying current return network;

FIG. 3 shows a cross section variation of electrically connecting theLSA gore to the underlying current return network with wire bonds;

FIG. 4 shows a cross section variation of a hole-type connection throughthe upper surface of the LSA gore to the underlying current returnnetwork.

DESCRIPTION OF PREFERRED EMBODIMENTS

Turning now to the drawings and more particularly, FIGS. 1A-B show anexample of an aircraft 100 protected by conductor 120 with lightningstrike appliqué (LSA) gores 102, 104 electrically attached and connectedthrough a flexible, electrical connection 106, according to anembodiment of the present invention. FIG. 1B is an expanded view of area101 of FIG. 1A. Suitable LSA gores 102, 104 are described in PublishedU.S. patent application Ser. No. 10/941,429, filed Sep. 15, 2004,Publication No. 2005/0181163, entitled “Appliqué,” published Aug. 18,2005, to Diane C. Rawlings et al. (Rawlings I) and assigned to theassigned of the present invention. Furthermore, LSA gores 102, 104 mayinclude a region of a Wide Area Lightning Diverter Overlay (WALDO). Anexample of a suitable WALDO is described in Published U.S. patentapplication Ser. No. 11/229,911, filed Sep. 19, 2005, Publication No.2006/051592, entitled “Wide Area Lightning Diverter Overlay,” publishedMar. 9, 2006, to Diane C. Rawlings et al. (Rawlings II) and assigned tothe assigned of the present invention. Regardless, however, preferredelectrical connects 106 provide a reliable electrical connection betweenthe underlying conductor 120 and LSA gores 102,104. Moreover, incombination with WALDO, the present invention provides a predictabledistribution path for current that may result from a lightning strike.

Both the conductor 120 and the LSA gores 102, 104 in FIGS. 1A-B aremanufactured typically as flexible flat multilayer laminates that areand readily bent and elongated typically in the range of 2-20% for easyapplication to a curved structure, e.g., an aircraft. Further, in thisexample, both the conductor 120 and LSA gores 102, 104 are adhesivelyattached to the aircraft composite skin. LSA gores 102, 104 typicallyinclude a polymer topcoat layer for protecting the conductive layer fromenvironmental elements and that, optionally, may be painted. Theflexible electrical connect 106 (preferably, contained subsurface toprotect the connect 106 from environmental damage) may be one or modewire bond loops, or one or more fuzz buttons inserted and connecting thepieces 104 to 102 and 120. Alternately, contact may be from above,passed through holes in the upper piece 104 to the lower 102, 120 orwith conductive adhesive bridging the conductive layers. In thisexample, the LSA gore 104 conducts current, e.g., from a lightningstrike, through the electrical connections 106 to the adjoining LSA andto the underlying conductor 120, which directs or dissipates currentfrom the lightning strike discharge.

The electrical connection 106 may be made between two LSA gores, oneoverlapping the other, and between an LSA gore and an underlying groundpath. Further, the size and shape of the LSA gores 102, 104, andunderlying conductor 120 are determined by surface curvature or tominimize aerodynamic drag or for ease of installation, e.g., LSA gores102, 104 on a wing may be trapezoidal. The conductor 120 may also besized as an integral part of the design of the lightning protectionsystem. Additionally, since the LSA gores 102, 104 may replace paint foreconomical static charge and lightning strike protection, the LSA gores102, 104 may cover only a part or substantially all of the compositeskin of aircraft 100.

Advantageously, a conductive LSA layer covering the composite skin ofthe aircraft 100 provides additional electromagnetic interference (EMI)shielding. For metal or composite skinned aircraft the LSA layer withelectrical connects also provides EMI shielding preventing radiationleaking through holes, gaps, and joints. EMI shielding may also beprovided for windows where the patterning of the foil in WALDO can forman inductive grid which enables the appliqué to be visibly transparentwhile remaining DC-conductive and connectable to LSA or other electricalground.

FIG. 2A shows a cross sectional example, wherein LSA gores 102, 104,both adhesively attached to composite skin 110, are electricallyconnected in a region of the overlap through A-A in FIG. 1B andsubstantially as described in Rawlings I. The LSA gores 102, 104 includea conductive layer 102C, 104C sandwiched in between a dielectric layer102D, 104D and a topcoat 102T, 104T. Each LSA gore 102, 104 alsoincludes an adhesive layer 102A, 104A, attaching the respective to theunderlying composite skin 110 and, with LSA gore 104, LSA gore 102. Thefirst LSA gore 102 is applied and attached directly to the compositeskin 110. An area 112 is removed from the topcoat 102T of the first LSAgore 102, exposing the conductive layer 102C.

Preferably, the area 112 along the edge of the second LSA gore 104(preferably 5-6 mm) from the edge is a small diameter annulus (<2-5 mm),where the dielectric layer 104D and the adhesive layer 104A have beenremoved leaving a small central region intact to hold a fuzz buttonconnector. Vias 114 are large enough to allow one end of the electricalconnect 116 (a fuzz button in this example) to contact the conductivelayer 104C and may be annular or single circular holes. Similarly, aregion in the topcoat of the first LSA gore 102T sized to accommodateready contact of the fuzz button (preferably 7-10 mm in length andwidth) or is removed. Preferably, these vias are created using a laserscribed template, but may be created more arduously using standardchemical etch or mechanical means. Normally, fuzz buttons are used insemiconductor test sockets and interconnects where low-distortiontransmission lines are a necessity. A typical fuzz button is fashionedfrom a single strand of wire, e.g., gold-plated beryllium copper wire.

In this example, the first LSA gore 102 is applied, e.g., pressed inplace, and adhesively fixed to, the composite skin 110. The fuzz buttonconnect 116, which is cylindrical but readily bent or doubled over intoa U-shape as in this example, may be inserted in the second LSA gore 104in vias 114 and, temporarily held in place by the adhesive layer 104A.Then, the second LSA gore 104 with the fuzz button 116 in place isapplied to the composite skin 110 to overlap 118 the first LSA gore 102,preferably by 10-15 mm. Preferably, the extent of overlap 118 isminimized to minimize weight and cost, but also for aesthetics and insome cases to facilitate adhesion of the LSA gores 104. This mayfacilitate adhesion because, depending on the topcoat composition andtexture, the adhesive may not adhere to the topcoat as well as to theaircraft surface. With the U-shaped fuzz button 116 held in place at theoverlap 118 by the adhesively attached second LSA gore 104, the fuzzbutton 116 mechanically contacts both conductive layers 102C, 104C for amuch improved electrical contact between the two.

FIG. 2B shows a cross section of a current return or ground connectionlayer 120 overlapped by and connected to LSA gore 104, e.g., through B-Bin FIG. 1B. In this example, the LSA gore 104 is connected to a currentreturn network 120 that is adhesively applied directly to the compositeskin 110. The current return network 120 includes a conductive surfacelayer 120C on a dielectric isolation layer 120D and is attached to thecomposite skin 110 by an adhesive layer 120A at the dielectric isolationlayer 120D. Typically the LSA gores 102, 104 and current return network120 are manufactured as flat, flexible multilayer laminates of aconductor layer 102C, 104C, 120C, preferably a metal foil or mesh, on adielectric film layer 102D, 104D, 120D of a suitable dielectric polymerand an adhesive attachment layer 102A, 104A, 120A. Unlike the LSA gores102, 104, the current return network 120 typically does not have atopcoat 102T, 104T; instead, the conductive surface layer 120C isexposed.

FIG. 3 shows a variation on the cross section of the ground connectionlayer overlapped by and connected to LSA gore 104 of FIG. 2B with likeelements labeled identically. In this example, instead of a fuzz button,wires 130 are wired bonded to connect a location of the foil on LSA gore104 to another location of the foil on LSA gore 104 and the wire loopmakes contact with the underlying conductor 120C. Preferably, the wires130 are 0.0005-0.001″ (12.5-25 μm) in diameter and operate substantiallythe same as the fuzz buttons 116 with U-shaped wires bonded in the vias114 to the second conductive layer 104C. When pressed in place, thewires 120 mechanically contact the first conductive layer 104,substantially the same as the fuzz buttons in the embodiment of FIG. 2A.

FIG. 4 shows another variation on the cross section of the groundconnection layer overlapped by and connected to LSA gore 104 of FIG. 2Bwith like elements labeled identically. Although not DC-conductive, thisconnector is particularly effective for transitioning the current from alightning strike traveling in a conductor such as LSA to the underlyingconductor of the current return network. In this example, instead offorming subsurface contacts between the LSA gores 102, 104, holes orthrough-vias 140 are bored through top of the second LSA gore 104 to theconductive layer 120C, or for a LSA gore conductive layer, exposing theconductive layer in that LSA gore. Further, for a LSA gore conductivelayer, an area of the topcoat of the LSA gore conductive layer may ormay not be removed (substantially as described for area 112 in FIG. 2A)prior to attaching the second LSA gore 104. The through-vias 140 areopened to the LSA gore conductive layer at these areas 112. Afteropening through-vias 140 in LSA gore 104, a fuzz button or other connectmay be inserted in the opened through-vias 140. For lightning protectionthe hole itself is an adequate conductor. For dissipatingPrecipitation-Static (P-Static) charge, for example, the through-vias140 may be filled with a suitable electrical conductor such as anelectrically conductive adhesive or sealant. Additionally, the holes maybe covered with a suitable dielectric protective seal if desired.

Advantageously, substantially the entire surface area of a compositestructure, such as a composite aircraft, may be covered with appliqués.The conductive layers and/or lightning diverter overlays distribute anddissipate current, such as from lightning strikes and thus, eliminateor, substantially mitigate any damage caused by lightning strikes. Theinterconnected appliqués are easy to install initially and, easy torepair. Because the surface applied appliqués, where applied, provide asubstantially uniform conductive layer, most of the current from eachlightning strike travels through the conductive appliqués rather thandamaging the underlying composite.

In addition, the interconnected appliqués conductive layers and/orlightning diverter overlays can cover the entire surface or any selectedportion thereof. The interconnected appliqués are relativelyinexpensive, with lightweight electrically connected appliquéseconomically providing lightning and exterior static charge protection.Also, preferred appliqués may be used as a discharge straps betweenareas where static charge is known to collect, e.g., windows in thevicinity of high radio frequency (RF) transmissions. So, the electricalconnects facilitate providing a deterministic scheme directing currentfrom a lightning strike, even in the absence of a DC connection, e.g.,appliqués with conductive through-vias. P-static charge can bedistributed and dissipated through appliqué-to-appliqué orpart-to-appliqué-to-part DC electrical connections.

Furthermore, interconnecting appliqués in a lightning protection systemaccording to a preferred embodiment of the present invention simplifiesdesign for static electricity discharge protection. The electricalconnects are flexible and thermally expandable to maintain connection ina lightning protection system and without excessively increasingaircraft weight but with superior performance and protection. Becauseappliqués are electrically connected a low resistance continuous currentpath is provided to minimize static charge build up, e.g., on theexterior of an aircraft. Thus, static-originated current passes acrossappliqué gores, joints or seams. Further, preferred electricalconnectors facilitate controlling the transition of lightning current(preferably, external to the skin of the aircraft), diverting currentfrom appliqué gores to an underlying current return network.

Thus, advantageously, high voltage effects and current arising from alightning strike may be prevented (or at least significantly attenuated)from penetrating the composite skin. Moreover, preferred electricalconnectors and especially subsurface connectors are readily located inany desired location without impacting the visual aesthetics orenvironmental durability of the surface. Finally, the preferredelectrical connectors and LSAs provide low cost and low weightelectromagnetic interference (EMI) shielding for aircraft or containmentboxes/vessels for sensitive electronics.

While the invention has been described in terms of preferredembodiments, those skilled in the art will recognize that the inventioncan be practiced with modification within the spirit and scope of theappended claims. It is intended that all such variations andmodifications fall within the scope of the appended claims. Examples anddrawings are, accordingly, to be regarded as illustrative rather thanrestrictive.

1. A method for preventing static charge buildup comprising: enablingcharge to distribute across a first conductive appliqué affixed to asurface; passing charge to a conductor coupled to said first conductiveappliqué; and directing the charge to a subsurface collection point. 2.A method for preventing static charge buildup as in claim 1, whereinsaid conductor is another conductive appliqué in a series of conductiveappliqués including said first conductive appliqué, charge beingdistributed over said series of conductive appliqués.
 3. A method forpreventing static charge buildup as in claim 2, wherein said firstconductive appliqué overlaps a portion of one or more other conductiveappliqués, charge being distributed among said series of conductiveappliqués at appliqué overlaps.
 4. A method for preventing static chargebuildup as in claim 3, further comprising passing charge from said firstconductive appliqué to said other conductive appliqués through orificesthrough non-conductive layers to conductive layers in each conductiveappliqué.
 5. A method for preventing static charge buildup as in claim4, wherein passing charge comprises passing current through conductorsin said orifices attached to the overlapping said conductive appliqués.6. A method for preventing static charge buildup as in claim 5, whereinsaid conductors are contained subsurface to said overlapping conductiveappliqués and are selected from the group consisting of a fuzz buttonand one or more wire bonds.
 7. A method for preventing static chargebuildup as in claim 2, wherein said structure is a composite materialstructure.
 8. A method for preventing static charge buildup as in claim7, wherein said structure is a composite aircraft.
 9. A method forpreventing static charge buildup as in claim 8, wherein said series ofconductive appliqués provide electromagnetic interference (EMI)shielding to said composite aircraft.
 10. A method for preventing staticcharge buildup as in claim 1, wherein said structure is an aircraft andsaid first conductive appliqué is a conductive strap passing charge overwindows.
 11. A method for preventing static charge buildup as in claim1, wherein said structure is an aircraft and said first conductiveappliqué is one of a plurality of conductive appliqués, each being aconductive strap passing charge over a respective aircraft seam orjoint.
 12. A method for preventing static charge buildup on an aircraft,said method comprising: distributing charge across each of a series ofconductive appliqués affixed to a surface, one or more conductiveappliqués overlapping other adjacent conductive appliqués; passingcharge vertically between said conductive appliqués at appliquéoverlaps; and directing accumulating charge from said series ofconductive appliqués to a subsurface collection point.
 13. A method forpreventing static charge buildup on an aircraft as in claim 12, whereinsaid aircraft is a composite aircraft.
 14. A method for preventingstatic charge buildup on a composite aircraft as in claim 13, whereinsaid series of conductive appliqués provide electromagnetic interference(EMI) shielding to said composite aircraft.
 15. A method for preventingstatic charge buildup on a composite aircraft as in claim 14, saidmethod further comprising passing charge from overlapping saidconductive appliqués to said adjacent conductive appliqués throughorifices through non-conductive layers to conductive layers in eachconductive appliqué.
 16. A method for preventing static charge buildupon a composite aircraft as in claim 15, wherein passing charge comprisespassing current through conductors in said orifices attached to saidoverlapping conductive appliqués.
 17. A method for preventing staticcharge buildup on a composite aircraft as in claim 16, wherein saidconductors are contained subsurface to said overlapping conductiveappliqués and are selected from the group consisting of a fuzz buttonand one or more wire bonds.
 18. A method for preventing static chargebuildup on an aircraft as in claim 13, further comprising passingcurrent through conductive appliqués affixed over windows.
 19. A methodfor preventing static charge buildup locally on an aircraft, said methodcomprising: distributing charge across the skin of an aircraft, chargepassing through conductive appliqués affixed over seams and joints;passing charge vertically at conductive appliqué edges; and directingaccumulating charge to a collection point.
 20. A method for preventingstatic charge buildup on an aircraft as in claim 19, wherein saidaircraft skin is conductive, passing charge vertically comprises passingcharge over conductors through orifices through non-conductive layers ineach conductive appliqué to said aircraft skin.